The invention described herein may be manufactured and used by or for the government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.
This invention relates to solid propellant formulations, which exhibit good processing properties, good safety characteristics, desirable combustion properties, good safety characteristics and excellent mechanical attributes. More specifically, this invention relates to solid propellant formulations, which have a burning rate of 0.5 in/s at 1000 psi and show no pressure slope break up to 8000 psi.
Solid propellants are used extensively in the aerospace industry. Solid propellants have developed as the preferred method of powering most missiles and rockets for military, commercial, and space applications. Solid rocket motor propellants have become widely accepted because of the fact that they are relatively simple to formulate and use, and they have excellent performance characteristics. Furthermore, solid propellant rocket motors are generally very simple when compared to liquid fuel rocket motors. For all of these reasons, it is found that solid rocket propellants are often preferred over other alternatives, such as liquid propellant rocket motors.
Conventional solid propellant rocket motors operate by generating large amounts of hot gases from the combustion of a solid propellant formulation stored in the motor casing. Typical solid rocket motor propellants are generally formulated having an oxidizing agent, a fuel, and a binder. At times, the binder and the fuel may be the same. In addition to the basic components set forth above, it is conventional to add various plasticizers, curing agents, cure catalysts, and other similar materials which aid in the processing and curing of the propellant. A significant body of technology has developed related solely to the processing and curing of solid propellants, and this technology is well known to those skilled in the art.
During operation, the gases generated from the combustion of the solid propellant accumulate within the combustion chamber until enough pressure is amassed within the casing to force the gases out of the casing and through an exhaust port. The expulsion of the gases from the rocket motor and into the environment produces thrust. Thrust is measured as the product of the total mass flow rate of the combustion products exiting the rocket multiplied by the velocity of the exiting combustion products plus the product of the change in pressure at the exit plane multiplied by the exit area. Increasing the pressure at which the gases are expelled from the combustion chamber raises the thrust level, which in turn increases the propulsion rate of the vehicle containing the rocket motor to thereby permit the vehicle to achieve higher speeds.
One type of propellant that is widely used incorporates ammonium perchlorate (AP) as the oxidizer. The AP oxidizer may then, for example, be incorporated into a propellant that is bound together by a hydroxy-terminated polybutadiene (HTPB) binder. Such binders are widely used and commercially available. It has been found that such propellant compositions provide ease of manufacture, relative ease of handling, good performance characteristics; and are at the same time economical and reliable. In essence it can be said that AP composite propellants have been the backbone of the solid propulsion industry for approximately the past 40 years.
High-performance booster propellants are being developed not only because they can deliver superior results but also because they can be used in numerous propulsion systems. One of the problems associated with conventional propellant formulations having an exponentially increasing propellant burn rate is that an increase consumption of propellant generally increases the operating pressure, which in turn increases the risk of catastrophic failure of the rocket motor casing. This insufficiency springs from the inherent combustion characteristics of AP, which often impart the propellant a high-pressure slope break between 2500 to 3500 psi. Therefore, to avoid this slope break region, propulsion systems designers generally devise motors that operate below 2000 psi, even though it is known that much higher performance can be achieved if the motor can operate at pressures higher than 2000 psi. The change in burning rate (rb) as a function of the pressure change, Pc, is defined as the burning rate slope, n:   N  =            (                                    ∂            ln                    ⁢                      xe2x80x83                    ⁢                      r            b                                                ∂            ln                    ⁢                      xe2x80x83                    ⁢                      P            c                              )        T  
Data for determining burning rates at different pressures are typically gathered either by standard strand testing or by test motor analysis. The determination of burning rates by such testing procedures is well known in the art. Generally, conventional solid rocket propellant formulations have burn rate slopes of 0.15 or greater, but well below 1.0. Conventional propellants usually exhibit a dramatic positive increase in burning rate slope at pressures above about 2500 psi to 3500 psi, as illustrated in FIG. 1.
The conventional solution to avoiding catastrophic failure of the rocket motor casing is to strengthen the rocket motor casing by constructing the casings with thick walls from strong, dense materials, such as steel. This approach, however, deleteriously imparts a severe weight penalty to the vehicle. Consequently, a greater amount of thrust and an increased propellant burn rate is required to propel the vehicle at a comparable rate. Because of the recent successes in manufacturing new composite motor cases, which can sustain much higher chamber pressures, it is possible to develop a motor that can deliver increased performance as measured by specific impulse (Isp).
Regardless of advances in rocket motor casings, it is highly desirable to search for a high-energy propellant material that does not exhibit a pressure slope break. While the exact cause is unknown, one theory is that it is due to a change in the mechanism from diffusion flame control to solid AP flame control at the higher pressures that exist in burning propellants, as described in xe2x80x9cA Review of Models and Mechanisms for Pressure Exponent Breaks in Composite Solid Propellantsxe2x80x9d, Cohen, N., Proceedings of 23rd JANNAF Combustion Meeting, CPIA Publication 457, Vol. II (October 1986), incorporated herein by reference. For example, burning solid AP is much more sensitive to pressure fluctuation (i.e., high slope) than materials experiencing diffusion-control-type burning.
Because conventional booster propellants (HTPB/AP/aluminum) are often composed of a mixture of different AP particle sizes, optimum packing efficiency (i.e., good processing properties) is important in attaining high performance and, at times, in providing burn rate tailorability. The various AP powders often contain large pieces (i.e., 200 xcexcm to 400 xcexcm) and it is extremely likely that the large-size AP particles contribute to the pressure slope break problem. Previous work has demonstrated that the slope break is delayed from 2500 psi to much higher pressures (as high as 10,000 psi) when the propellant formulations do not contain large AP particles. Therefore, many formulations instead incorporate high levels of ultrafine AP and exotic burn rate catalysts, such as superfine iron oxide, chromic oxide, catocene, or carboranes to achieve the higher pressure slope break.
Attempts have been made to solve the problems associated with the high-pressure slope break that occurs between 2500 psi and 3500 psi with ammonium perchlorate as an oxidizer. The problem is discussed in detail in the following publications: Atwood, A. I. et al., xe2x80x9cHigh-pressure Burning Rate Studies of Ammonium Perchlorate (AP) Based Propellants,xe2x80x9d Proceedings for Research and Technology Agency of North Atlantic Treaty Organization (NATO) 1999 Meeting on Small Rocket Motors and Gas Generators for Land, Sea, and Air Launched Weapon Systems, 19-23 April 1999, Corfu, Greece and Boggs, T. L., et al., xe2x80x9cAmmonium Perchlorate Combustion: Effects of Sample Preparation; Ingredient Type; and Pressure, Temperature and Acceleration Environments,xe2x80x9d J Combustion Science and Technology, Vol. 7, pp. 177-183 (1973), incorporated herein by reference. U.S. Pat. No. 6,086,692 issued to Hawkins et al. on Jul. 11, 2000 discloses a solid rocket propellant formulation operable at high pressures with burn rates relatively insensitive to changes in pressure and propellant temperature. However, the propellant formulation of U.S. Pat. No. 6,086,692 shows a double plateau, rather than the constant low slope of the invention of the current application. Also, the propellant formulation of U.S. Pat. No. 6,086,692 does not combine a good burning rate with good hazard properties like the invention of the current application. In addition the invention of the current application uses different sets of ingredients and particle sizes to achieve a performance not found in the propellant formulation of U.S. Pat. No. 6,086,692.
It is an object of the present invention to produce a IMAD-213 AP solid propellant formulation having a pressure slope break higher than about 2500 psi to 3500 psi comprising about 6.0 to about 6.5 weight % of at least one energetic polymeric binder; about 25 to about 32 weight % ammonium perchlorate having a particle size of about 10 xcexcm to about 15 xcexcm as a primary oxidizer; about 18 to about 24 weight % ammonium perchlorate having a particle size of about 80 xcexcm to about 100 xcexcm as a secondary oxidizer; about 10 to about 15 weight % of ammonium nitrate having a particle size of about 40 xcexcm to about 60 xcexcm as a co-oxidizer; about 19 to about 24 weight % of a metal fuel; and about 0.3 to about 0.6 weight % of a burn rate catalyst. Further, the solid propellant formulation bums at a rate of about 0.5 in/s at about 1000 psi and shows no pressure slope break up to about 8000 psi. Also, the IMAD-213 AP solid propellant formulation may contain a plasticizer, a curative, a stabilizer, a cure catalyst, and a bonding agent, depending upon the desired characteristics of the propellant.
Another object of the present invention is to provide a IMAD-213 CL-20 solid propellant formulation having a pressure slope break occurring higher than about 2500 psi to 3500 psi comprising about 6.0 to about 6.5 weight % of at least one energetic polymeric binder; about 18 to about 24 weight % ammonium perchlorate having a particle size of about 80 xcexcm to about 100 xcexcm as a primary oxidizer; about 10 to about 15 weight % of CL-20 having a particle size less than about 2.0 xcexcm as a secondary oxidizer; about 25 to about 32 weight % of ammonium nitrate having a particle size of about 40 xcexcm to about 60 xcexcm as a co-oxidizer; about 19 to about 24 weight % of a metal fuel; and about 0.3 to about 0.6 weight % of a burn rate catalyst. Further, the solid propellant formulation burns at a rate of about 0.5 in/s at about 1000 psi and shows no pressure slope break up to about 8000 psi. The IMAD-213 CL-20 solid propellant formulation may contain a plasticizer, a curative, a stabilizer, a cure catalyst, and a bonding agent, depending upon the desired characteristics of the propellant.
Another objective of a preferred embodiment of the present invention is to provide a IMAD-213 RDX solid propellant formulation having a pressure slope break higher than about 2500 psi to 3500 psi comprising: about 6.0 to about 6.5 weight % of at least one energetic polymeric binder; about 18 to about 22 weight % ammonium perchlorate having a particle size of about 10 xcexcm to about 15 xcexcm as a first oxidizer; about 18 to about 22 weight % ammonium perchlorate having a particle size of about 80 xcexcm to about 100 xcexcm as a second oxidizer; about 8 to about 10 weight % of RDX having a particle size of about 1.7 xcexcm to about 2.5 xcexcm as a third oxidizer; about 10 to about 15 weight % of ammonium nitrate having a particle size of about 40 xcexcm to about 60 xcexcm as a co-oxidizer; about 20 to about 24 weight % of a metal fuel; and about 0.2 to about 0.5 weight % of a burn rate catalyst. Further, the solid propellant formulation burns at a rate of about 0.5 in/s at about 1000 psi and shows no pressure slope break up to about 8000 psi. The IMAD-213 RDX solid propellant formulation may contain a plasticizer, a curative, a stabilizer, a cure catalyst, and a bonding agent, depending upon the desired characteristics of the propellant.
One object of the present invention is to provide high-performance shock insensitive booster propellant formulations which burn at a constant rate of about 0.5 in/s at about 1000 psi and do not exhibit a pressure slope break between about 2500 to 3500 psi.
Another object of the invention is to provide high-performance shock insensitive booster propellant formulations in which the pressure slope break is delayed to pressures as high as 10,000 psi.
Another object of the invention is to provide high-performance shock insensitive booster propellant formulations that avoid the use of AP having a particle size from about 200 xcexcm to about 400 xcexcm.
A still further object of the invention is to provide high-performance shock insensitive booster propellant formulations which are environmentally friendly and do not incorporate exotic burn rate catalysts such as superfine iron oxide, chromic oxide, catocene, or carboranes.